Variable fan nozzle using shape memory material

ABSTRACT

A gas turbine engine ( 10 ) includes a fan ( 14 ), a nacelle ( 28 ) arranged about the fan, and an engine core at least partially within the nacelle. A fan bypass passage ( 30 ) downstream of the fan between the nacelle and the gas turbine engine conveys a bypass airflow ( 1 ) from the fan. A nozzle ( 40 ) associated with the fan bypass passage is operative to control the bypass airflow. The nozzle includes a shape memory material having a first solid state phase that corresponds to a first nozzle position and a second solid state phase that corresponds to a second nozzle position.

BACKGROUND OF THE INVENTION

This invention relates to gas turbine engines and, more particularly, toa gas turbine engine having a variable fan nozzle for controlling abypass airflow through a fan bypass passage of the gas turbine engine.

Gas turbine engines are widely known and used for power generation andvehicle (e.g., aircraft) propulsion. A typical gas turbine engineincludes a compression section, a combustion section, and a turbinesection that utilize a core airflow into the engine to propel thevehicle. The gas turbine engine is typically mounted within a housing,such as a nacelle. A bypass airflow flows through a passage between thehousing and the engine and exits from the engine at an outlet.

Presently, conventional gas turbine engines are designed to operatewithin a desired performance envelope under certain predetermined flightconditions, such as cruise. Conventional engines tend to approach orexceed the boundaries of the desired performance envelope under flightconditions outside of cruise, such as take-off and landing, which mayundesirably lead to less efficient engine operation. For example, thesize of the fan and the ratio of the bypass airflow to the core airfloware designed to maintain a desired pressure ratio across the fan duringcruise. However, during take-off and landing, the pressure ratio maychange such that pressure pulsations occur across the fan (i.e., fanflutter). The pressure pulsations cause less efficient fan operation andincrease mechanical stress on the fan, which ultimately causes anincrease in fuel consumption and reduces the life expectancy of the fan.

Therefore, there is a need to control the bypass airflow over a widervariety of different flight conditions, such as take-off and lift-off,to enable enhanced control of engine operation.

SUMMARY OF THE INVENTION

An example gas turbine engine includes a fan, a nacelle arranged aboutthe fan, and an engine core at least partially within the nacelle. A fanbypass passage downstream of the fan between the nacelle and the gasturbine engine conveys a bypass airflow from the fan. A nozzleassociated with the fan bypass passage is operative to control thebypass airflow. The nozzle includes a shape memory material having afirst solid state phase that corresponds to a first nozzle position anda second solid state phase that corresponds to a second nozzle position.The shape memory material is thermally active. A controller controls anactuator near the nozzle selectively heats and cools the nozzle toreversibly transition the shape memory material between the phases tomove the nozzle between the positions.

In one example, the nozzle includes triangular-shaped tabs that extendfrom the nacelle. The controller selectively axially extends andretracts the tabs to change an effective area of the nozzle. In anotherexample, the controller selectively moves the tabs in a radial directionto change the effective area.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows.

FIG. 1 illustrates selected portions of an example gas turbine enginesystem with a nozzle having shape memory material.

FIG. 2 schematically illustrates an example tab of the nozzle, whereinthe tab is axially moveable.

FIG. 3 schematically illustrates an example tab of the nozzle, whereinthe tab is radially moveable.

FIG. 4 schematically illustrates an example tab of the nozzle, having awire frame made of shape memory material.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 illustrates a schematic view of selected portions of an examplegas turbine engine 10 suspended from an engine pylon 12 of an aircraft,as is typical of an aircraft designed for subsonic operation. The gasturbine engine 10 is circumferentially disposed about an enginecenterline, or axial centerline axis A. The gas turbine engine 10includes a fan 14, a low pressure compressor 16 a, a high pressurecompressor 16 b, a combustion section 18, a low pressure turbine 20 a,and a high pressure turbine 20 b. As is well known in the art, aircompressed in the compressors 16 a, 16 b is mixed with fuel that isburned in the combustion section 18 and expanded in the turbines 20 aand 20 b. The turbines 20 a and 20 b are coupled for rotation with,respectively, rotors 22 a and 22 b (e.g., spools) to rotationally drivethe compressors 16 a, 16 b and the fan 14 in response to the expansion.In this example, the rotor 22 a also drives the fan 14 through a geartrain 24.

In the example shown, the gas turbine engine 10 is a high bypassturbofan arrangement. In one example, the bypass ratio is greater than10, and the fan 14 diameter is substantially larger than the diameter ofthe low pressure compressor 16 a. The low pressure turbine 20 a has apressure ratio that is greater than 5, in one example. The gear train 24can be any known suitable gear system, such as a planetary gear systemwith orbiting planet gears, planetary system with non-orbiting planetgears, or other type of gear system. In the disclosed example, the geartrain 24 has a constant gear ratio. Given this description, one ofordinary skill in the art will recognize that the above parameters areonly exemplary and that other parameters may be used to meet theparticular needs of an implementation.

An outer housing, nacelle 28, (also commonly referred to as a fannacelle or outer cowl) extends circumferentially about the fan 14. Agenerally annular fan bypass passage 30 extends between the nacelle 28and an inner housing, inner cowl 34, which generally surrounds thecompressors 16 a, 16 b and turbines 20 a, 20 b.

In operation, the fan 14 draws air into the gas turbine engine 10 as acore flow, C, and into the bypass passage 30 as a bypass air flow, D. Inone example, approximately 80 percent of the airflow entering thenacelle 28 becomes bypass airflow D. A rear exhaust 36 discharges thebypass air flow D from the gas turbine engine 10. The core flow C isdischarged from a passage between the inner cowl 34 and a tail cone 38.A significant amount of thrust may be provided by the discharge flow dueto the high bypass ratio.

The example gas turbine engine 10 shown FIG. 1 also includes a nozzle 40(shown schematically) associated with the bypass passage 30. In thisexample, the nozzle 40 is shown near the rear of the nacelle 28,however, in other examples, the nozzle 40 is located farther forward butaft of the fan 14. In this example, the nozzle 40 is coupled to thenacelle 28. Alternatively, the nozzle 40 is coupled with the inner cowl34, or other suitable structural portion of the gas turbine engine 10.

The nozzle 40 is operative to move between a plurality of positions toinfluence the bypass air flow D, such as to manipulate an air pressureof the bypass air flow D. A controller 44 commands the nozzle 40 toselectively move among the plurality of positions to manipulate thebypass air flow D in a desired manner. The controller 44 may bededicated to controlling the nozzle 40, integrated into an existingengine controller within the gas turbine engine 10, or be incorporatedwith other known aircraft or engine controls. For example, selectivemovement of the nozzle 40 permits the controller 44 to vary the amountand direction of thrust provided, enhance conditions for aircraftcontrol, enhance conditions for operation of the fan 14, or enhanceconditions for operation of other components associated with the bypasspassage 30, depending on input parameters into the controller 44.

In one example, the gas turbine engine 10 is designed to operate withina desired performance envelope under certain predetermined conditions,such as cruise. For example, the fan 14 is designed for a particularflight condition—typically cruise at 0.8 Mach and 35,000 feet. The fan14 is designed at a particular fixed stagger angle for an efficientcruise condition. The nozzle 40 is operated to influence the bypassairflow D such that the angle of attack or incidence on the fan 14 ismaintained close to design incidence at other flight conditions, such aslanding and takeoff, thus enabling a desired engine operation over arange of flight condition with respect to performance and otheroperational parameters such as noise levels. In one example, it isdesirable to operate the fan 14 under a desired pressure ratio range(i.e., the ratio of air pressure forward of the fan 14 to air pressureaft of the fan 14) to avoid fan flutter. To maintain this range, thenozzle 40 is used to influence the bypass airflow D to control the airpressure aft of the fan 14 and thereby control the pressure ratio. Insome examples, the nozzle 40 varies an effective area associated withthe nozzle 40 by approximately 20% to influence the bypass airflow D.

In the illustrated example, the nozzle 40 includes tabs 54 that extendin a generally axial direction from the nacelle 28. Referring to FIG. 2,each of the tabs 54 is generally triangular-shaped and tapers from theforward end to the trailing end. It is to be understood that thedesigned shape of the tabs 54 can vary from the illustrated example. Onebenefit associated with using the tabs 54 is noise reduction due tomixing of the exiting bypass airflow D with the surrounding outerairflow.

Each of the tabs 54 in this example is a hollow section that defines aninternal cavity 56. The internal cavity 56 tapers with the taper of thewalls of the tab 54. An actuator 58 is mounted within the internalcavity 56 in a known manner. Alternatively, the actuator 58 is mountedoutside of the internal cavity 56 within the nacelle 28 or other portionof the engine 10.

The tabs 54 are made of a shape memory material (“SMM”) that permits thetabs 54 to move between a plurality of different positions forcontrolling the effective area of the nozzle 40 to thereby influence thebypass airflow D. In one example, the SMM is thermally active andchanges shape in response to a change in temperature. In this example,the controller 44 selectively activates the actuator 58 to move the tabs54 using radiant heat. The tabs 54 move in response to the amount ofheat relative to a threshold temperature associated with the SMM. Bycontrolling the amount of heat, the controller 44 is able to control theposition of the tabs 54 to obtain a desired effective area.

In the illustrated example, the actuator 58 includes a control unit 60that is connected to the controller 44 and a wire 62 that forms a loopwithin the internal cavity 56. The control unit 60 is operative totransmit an electric current through the wires 62 to thereby resistivelyheat the wire 62 and produce heat that radiates throughout the internalcavity 56. The wire 62 alternatively includes multiple loops that form aheating mesh to uniformly heat the tab 54. Each tab 54 may includes itsown actuator 58. Alternatively, a single actuator 58 heats one or morewires 62 that extend through multiple internal cavities 56.

The controller 44 selectively actives the actuators 58 to move the tabs54. In a closed, extended position designated by the solid line of thetab 54 (FIG. 2), the nozzle 40 has an associated effective area, AR. Inan open, retracted position designated by the dashed line of the tab54′, the nozzle 40 has an associated effective area, AR′. Thus, in thisexample, moving the nozzle 40 between the closed and open positionchanges the effective area by an amount proportional to the differencebetween AR and AR′. In this example, the difference is proportional tothe change in axial length of the tab 54.

In another example, the tabs 54 move in a radial direction rather thanthe axial direction. In an open position designated by the solid line ofthe tab 54, the nozzle 40 has an associated effective area, AR. In aclosed position designated by the dashed line of the tab 54′, the nozzle40 has an associated effective area, AR′. Thus, in this example, movingthe nozzle 40 between the closed and open position changes the effectivearea by an amount approximately proportional to the change in radialdistance between the tab 54 and the inner cowl 34. As can beappreciated, the tab 54 may also move axially and radially for a changein the effective area.

In the above examples, the controller 44 selectively commands the tabs54 to move in order to obtain a desired effective area for controllingthe air pressure of the bypass airflow D within the bypass passage 30.For example, closing the nozzle 40 (AR in the above example) restrictsthe bypass airflow D and produces a pressure build-up (i.e., an increasein air pressure) within the bypass passage 30. Inversely, opening thenozzle 40 (AR′ in the above examples) permits more bypass airflow D andreduces the pressure build-up (i.e., a decrease in air pressure). Asdescribed above, controlling the bypass airflow provides numerousbenefits associated with the fan 14, engine thrust, etc.

In one example, the SMM is characterized as being reversibly changeablebetween a first solid state phase that corresponds to one of the nozzle40 positions and a second solid state phase that corresponds to anotherof the nozzle 40 positions. The phases represent differentcrystallographic arrangements of the atomic elements of the SMM. In oneexample, the SMM transitions between the phases relative to a thresholdtemperature (e.g., a temperature range in some examples) such that abovethe threshold temperature the SMM transitions into one of the phases andbelow the threshold temperature the SMM transitions into the otherphase.

The SMM is “trained” in a known manner by heating and deforming the SMMsuch that the two phases “remember” the shape of the tab 54 that isassociated with that phase. Thus, the controller 44 selectively heatsthe tabs 54 using the actuators 58 or cools the tabs 54 by shutting offthe actuators 58 (permitting the surrounding airflow to cool the tabs54) to transition the SMM between the phases and thereby move the tabs54 between the different positions.

FIG. 4 illustrates another embodiment of an example tab 54. In thisexample, the tab 54 is made from a non-continuous sheet made of SMMinstead of continuous sheets as in the above example. In the illustratedexample, the non-continuous sheet includes wire frame 74 having wires 76that are made of the SMM. The wire frame 74 supports a flexible skin 78that generally moves with the wires 76 between the different positionsto influence the bypass airflow D. Given this description, one ofordinary skill in the art will recognize suitable types of flexible skin78 to meet their particular needs.

In one example, the smart memory material of the above examples is anickel-titanium alloy, and the two phases are austenite and martensite.Other example thermally active SMM's include certain copper alloys,nickel alloys, cobalt alloys, manganese alloys, copper-aluminum alloys,copper-zinc-aluminum alloys, and combinations thereof. Generally, one ormore of the example thermally active SMM's are known as having shapememory characteristics over a temperature range of approximately −150°C. to 200° C. Additionally, temperature changes of between about 2° C.to 20° C. about the threshold temperature of one or more of the exampleSMM's are enough to transition between phases, which provides thebenefit of not having to heat or cool the tabs 54 over a largetemperature range to change nozzle 40 positions.

Although a preferred embodiment of this invention has been disclosed, aworker of ordinary skill in this art would recognize that certainmodifications would come within the scope of this invention. For thatreason, the following claims should be studied to determine the truescope and content of this invention.

What is claimed is:
 1. A gas turbine engine comprising: a fan disposedabout an engine axis; a nacelle arranged about the fan; an engine coreat least partially within the nacelle, the engine core having acompressor and a turbine; a fan bypass passage downstream of the fanbetween the nacelle and the engine core, for conveying a bypass airflowfrom the fan; a nozzle section disposed about the engine axis forcontrolling the bypass airflow, wherein the nozzle section includes ashape memory material disposed at least partially around an internalcavity and defining a radial inner side and a radial outer side of theinternal cavity, the shape memory material having a first solid statephase that corresponds to a first nozzle position and a second solidstate phase that corresponds to a second nozzle position; and anactuator at least partially within the internal cavity, the actuatorhaving a radiant heating element that includes wires for selectivelyheating the internal cavity and the shape memory material.
 2. The gasturbine engine recited in claim 1, wherein the shape memory materialcomprises a threshold temperature that corresponds to a reversiblechange between the first solid state phase and the second solid statephase.
 3. The gas turbine engine recited in claim 1, wherein the shapememory material comprises a nickel-titanium alloy.
 4. The gas turbineengine recited in claim 1, wherein the shape memory material comprises amaterial selected from a copper alloy, a nickel alloy, a cobalt alloy, amanganese alloy, a copper-aluminum alloy, a copper-zinc-aluminum alloy,and combinations thereof.
 5. The gas turbine engine of claim 1, whereinthe nozzle includes tabs that include the shape memory material.
 6. Thegas turbine engine recited in claim 5, wherein the tabs extend from thenacelle in a generally axial direction relative to an axis of rotationof the fan.
 7. The gas turbine engine recited in claim 5, wherein thetabs taper from a forward end toward a trailing end.
 8. The gas turbineengine recited in claim 1, wherein the nozzle section moves in an axialdirection between the first nozzle position and the second nozzleposition.
 9. The gas turbine engine recited in claim 1, wherein thenozzle section moves in a radial direction between the first nozzleposition and the second nozzle position.
 10. The gas turbine enginerecited in claim 1, wherein the shape memory material comprises a cobaltalloy.
 11. The gas turbine engine recited in claim 1, wherein the shapememory material consists of a cobalt alloy.
 12. The gas turbine enginerecited in claim 1, wherein the nozzle varies an effective areaassociated with the nozzle by approximately 20% to influence the bypassairflow.
 13. The gas turbine engine recited in claim 1, wherein a singleactuator heats the wires extending through multiple internal cavities.14. The gas turbine engine recited in claim 5, wherein the tabs are onlyshape memory material.
 15. The gas turbine engine recited in claim 5,wherein the internal cavity tapers with a taper of walls of the tabs.16. The gas turbine engine recited in claim 5, wherein the wiresincludes multiple loops that form a heating mesh to uniformly heat thetabs.
 17. The gas turbine engine recited in claim 5, wherein each tab ofthe tabs includes one of a plurality of actuators.
 18. A variable fannozzle for use in a gas turbine engine, comprising: a nozzle section forinfluencing a bypass airflow associated with a fan bypass passage, thenozzle section including a shape memory material disposed at leastpartially around an internal cavity and defining a radial inner side anda radial outer side of the internal cavity, the shape memory materialhaving a first solid state phase that corresponds to a first nozzleposition and a second solid state phase that corresponds to a secondnozzle position, and an actuator at least partially within the internalcavity, the actuator having a radiant heating element that includeswires for selectively heating the internal cavity and the shape memorymaterial.
 19. A method for controlling a bypass airflow through a nozzleassociated with a fan bypass passage in a gas turbine engine, comprisingthe steps of: controlling a radiant heating element within an internalcavity of the nozzle to provide a desired temperature within theinternal cavity; and controlling the desired temperature to selectivelyreversibly transition a shape memory material, disposed at leastpartially around the internal cavity of the nozzle and defining a radialinner side and a radial outer side of the internal cavity, relative to athreshold temperature associated with the shape memory material, wherethe reversible transitioning includes transitioning between a firstsolid state phase and a second solid state phase of the shape memorymaterial to move the nozzle between a first position corresponding tothe first solid state phase and a second position corresponding to thesecond solid state phase.
 20. The method recited in claim 19, furtherincluding reversibly transitioning between the first solid state phaseand the second solid state phase to change an effective area of thenozzle.
 21. The method recited in claim 19, further including reversiblytransitioning between the first solid state phase and the second solidstate phase to move the nozzle in an axial direction between the firstposition and the second position relative to a central axis of the gasturbine engine.
 22. The method recited in claim 19, further includingreversibly transitioning between the first solid state phase and thesecond solid state phase to move the nozzle in a radial directionbetween the first position and the second position relative to a centralaxis of the gas turbine engine.